U.S. patent application number 11/829358 was filed with the patent office on 2009-01-29 for in-situ brazing methods for repairing gas turbine engine components.
This patent application is currently assigned to HONEYWELL INTERNATIONAL, INC.. Invention is credited to ALBERT F. GONZALEZ, YIPING HU, CLYDE R. TAYLOR.
Application Number | 20090026182 11/829358 |
Document ID | / |
Family ID | 40294332 |
Filed Date | 2009-01-29 |
United States Patent
Application |
20090026182 |
Kind Code |
A1 |
HU; YIPING ; et al. |
January 29, 2009 |
IN-SITU BRAZING METHODS FOR REPAIRING GAS TURBINE ENGINE
COMPONENTS
Abstract
Methods are provided for repairing cracks in a damaged section
of a gas turbine engine component with an in-situ brazing process.
In an embodiment, by way of example only, the method includes
applying a braze paste to the damaged section of the component, the
braze paste comprising a braze material and an organic binder. The
method also includes subjecting the damaged section of the
component to a first temperature that is below a brazing
temperature of the braze material to thereby substantially
decompose and evaporate the organic binder, and heating the braze
material using laser energy to a second temperature that is
substantially equal to or above the brazing temperature to form the
brazed joint on the component.
Inventors: |
HU; YIPING; (GREER, SC)
; TAYLOR; CLYDE R.; (LAURENS, SC) ; GONZALEZ;
ALBERT F.; (GREENVILLE, SC) |
Correspondence
Address: |
HONEYWELL INTERNATIONAL INC.
101 COLUMBIA ROAD, P O BOX 2245
MORRISTOWN
NJ
07962-2245
US
|
Assignee: |
HONEYWELL INTERNATIONAL,
INC.
Morristown
NJ
|
Family ID: |
40294332 |
Appl. No.: |
11/829358 |
Filed: |
July 27, 2007 |
Current U.S.
Class: |
219/121.64 ;
219/72 |
Current CPC
Class: |
F01D 5/005 20130101;
Y02T 50/671 20130101; B23P 6/045 20130101; F05D 2230/237 20130101;
B23P 6/007 20130101; B23K 1/0018 20130101; B23K 2101/001 20180801;
Y02T 50/60 20130101; B23K 1/0056 20130101 |
Class at
Publication: |
219/121.64 ;
219/72 |
International
Class: |
B23K 26/20 20060101
B23K026/20; B23K 26/42 20060101 B23K026/42 |
Claims
1. An in-situ brazing method for repairing a crack in a damaged
section of a component, the method comprising the steps of:
applying a braze paste to the damaged section of the component, the
braze paste comprising a braze material and an organic binder;
subjecting the damaged section of the component to a first
temperature that is below a brazing temperature of the braze
material to thereby substantially decompose and evaporate the
organic binder; and heating the braze material with laser energy to
a second temperature that is substantially equal to or above the
brazing temperature to form a brazed joint on the component.
2. The method of claim 1, further comprising removing surface
contaminants from an area around the crack formed in the damaged
section of the component.
3. The method of claim 2, wherein the step of removing comprises:
chemically removing oxides from the component.
4. The method of claim 2, wherein the step of removing comprises:
mechanically removing oxides from the component.
5. The method of claim 4, wherein the step of removing further
comprises: applying acetone to clean the component.
6. The method of claim 1, wherein the step of subjecting the
damaged section of the component to a first temperature comprises
subjecting the component to a vacuum.
7. The method of claim 1, wherein the step of subjecting the
damaged section of the component to a first temperature comprises
exposing the component to a temperature in a range of about
500.degree. C. to about 550.degree. C. for a period of time in a
range of between about 0.5 and about 1 hour.
8. The method of claim 1, wherein the step of subjecting the
damaged section of the component to a first temperature comprises
heating a portion of the component on which the braze material is
applied to the first temperature with a laser.
9. The method of claim 1, wherein the step of heating the braze
material comprises exposing the component to an inert gas
atmosphere.
10. The method of claim 9, wherein the step of heating the braze
material comprises exposing the component to an argon
atmosphere.
11. The method of claim 1, further comprising machining the
component to an original shape and an original dimension.
12. An in-situ brazing method for repairing a crack in a damaged
section of a component comprising a superalloy to form a brazed
joint, the method comprising the steps of: removing surface
contaminants from an area around the crack formed in the damaged
section of the superalloy component; applying a braze paste to the
damaged section of the superalloy component, the braze paste
including a braze material and an organic binder; heating the
damaged section of the superalloy component in vacuum furnace to a
temperature in a range of between about 500 and about 550.degree.
C. for a period of time between about 0.5 and about 1 hour to
thereby substantially decompose and evaporate the organic binder;
and heating the braze material with laser energy to a second
temperature that is substantially equal to or above the brazing
temperature to form the brazed joint on the superalloy
component.
13. The method of claim 12, wherein the step of removing comprises:
chemically removing oxides from the superalloy component.
14. The method of claim 12, wherein the step of removing comprises:
mechanically removing oxides from the component.
15. The method of claim 14, wherein the step of removing further
comprises: applying acetone to clean the superalloy component.
16. The method of claim 12, wherein the step of subjecting the
damaged section of the superalloy component to a first temperature
comprises subjecting the superalloy component to a vacuum.
17. The method of claim 12, wherein the step of subjecting the
damaged section of the superalloy component to a first temperature
comprises heating a portion of the superalloy component on which
the braze material is applied to the first temperature with a
laser.
18. The method of claim 12, wherein the step of heating the braze
material comprises exposing the superalloy component to an inert
gas atmosphere.
19. The method of claim 18, wherein the step of heating the braze
material comprises exposing the superalloy component to an argon
atmosphere.
20. The method of claim 12, further comprising machining the
component to an original shape and an original dimension.
Description
TECHNICAL FIELD
[0001] The inventive subject matter generally relates to metallic
components of gas turbine engines, and more particularly relates to
methods for repairing turbine engine components.
BACKGROUND
[0002] Turbine engines are used as a primary power source for
various kinds of aircraft. Most turbine engines generally follow
the same basic power generation procedure. Air is ingested into a
fan section, and passes over stator vanes that direct the air into
a compressor section to be compressed. The compressed air is flowed
into a combustor, is mixed with fuel and burned, and the expanding
hot gases are directed, at a relatively high velocity, against
stationary turbine vanes in the engine. The vanes turn the high
velocity gas flow partially sideways to impinge on blades mounted
on a rotatable turbine disk. The force of the impinging gas causes
the turbine disk to spin at high speeds. Jet propulsion engines use
the power created by the rotating turbine disk to draw more air
into the engine and the high velocity gas is passed out of the
turbine to create forward thrust.
[0003] After repeated operation, some components may experience
thermal fatigue, oxidation and/or corrosion degradation. As a
result, the component may become damaged and may develop small
cracks and/or materials loss therein. To repair these components,
conventional welding techniques, such as plasma transferred arc
(PTA) welding or tungsten inert gas (TIG) welding, have been used
in the past. Typically for these techniques, the component is
placed in an inert gas atmosphere, and a filler material is then
welded to a damaged section of the component.
[0004] Although conventional welding techniques are useful for
repairing some components of the turbine engine, they have some
drawbacks when repairing others. For example, some components, such
as housings used in the combustor, air diffusers used in the
compressor, and bearing support housings, may be made of different
kinds of sheet metals. Thus, during welding operation when the
component is heated to high temperatures, it may be experience
different strain and stress levels in different areas due to
varying rates of deformation in those areas. Consequently, the
component may develop additional cracks, which may result in
repeated repairs, or discard and replacement of the component. In
another example, components made of two or more kinds of materials
having different thermal expansion coefficients may also develop
additional cracks, if subjected to PTA or TIG welding techniques.
Specifically, these techniques may cause hot cracking and/or part
distortion due to relatively excessive heat input.
[0005] Accordingly, an improved method for repairing cracks is
desired. In particular, it is desirable to have a method that does
not cause the formation of additional cracks in a component to be
repaired. In addition, it is desirable for the method to be
relatively simple and inexpensive to implement. Furthermore, other
desirable features and characteristics of the inventive subject
matter will become apparent from the subsequent detailed
description of the inventive subject matter and the appended
claims, taken in conjunction with the accompanying drawings and
this background of the inventive subject matter.
BRIEF SUMMARY
[0006] Methods are provided for repairing a damaged section of a
component. In an embodiment, by way of example only, the method
includes applying a braze paste to the damaged section of the
component, the braze paste comprising a braze material and an
organic binder. The method also includes subjecting the damaged
section of the component to a first temperature that is below a
brazing temperature of the braze material to thereby substantially
decompose and evaporate the organic binder, and heating the braze
material to a second temperature that is substantially equal to or
above the brazing temperature using laser energy.
[0007] In another embodiment, a method is provided for repairing a
crack or material loss in a damaged section of a component
comprising a superalloy with a brazed joint or buildup. Surface
contaminants are removed from an area around the crack formed in
the damaged section of the superalloy component. A braze paste is
applied to the damaged section of the superalloy component. The
braze paste includes a braze material and an organic binder. The
damaged section of the superalloy component is heated in a vacuum
furnace to a temperature in a range of between about 500.degree. C.
to 550.degree. C. for a period of time in a range of between about
0.5 to 1.0 hour to thereby substantially decompose and evaporate
the organic binder. The braze material is heated using laser energy
to a second temperature that is substantially above the brazing
temperature to form the brazed joint.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The inventive subject matter will hereinafter be described
in conjunction with the following drawing figures, wherein like
numerals denote like elements, and
[0009] FIG. 1 is a partial cross-sectional side view of a turbofan
jet engine, according to an embodiment;
[0010] FIG. 2 is a cross-sectional view of a damaged section of a
component including a crack, according to an embodiment; and
[0011] FIG. 3 is a flow diagram of a method for repairing
components of a turbofan jet engine, according to an
embodiment.
DETAILED DESCRIPTION
[0012] The following detailed description is merely exemplary in
nature and is not intended to limit the inventive subject matter or
the application and uses of the inventive subject matter.
Furthermore, there is no intention to be bound by any theory
presented in the preceding background or the following detailed
description.
[0013] Turning now to the description and with reference first to
FIG. 1, a partial cross-sectional side view of a turbofan jet
engine 100 is depicted. The turbofan jet engine 100 includes a fan
module 110, a compressor module 120, a combustor and turbine module
130, and an exhaust module 140. The fan module 110 is positioned at
the front, or "inlet" section of the engine 100, and includes a fan
108 that induces air from the surrounding environment into the
engine 100. The fan module 110 accelerates a fraction of this air
toward the compressor module 120, and the remaining fraction is
accelerated into and through a bypass 112, and out the exhaust
module 140. The compressor module 120 raises the pressure of the
air it receives to a relatively high level.
[0014] The high-pressure compressed air then enters the combustor
and turbine module 130, where a ring of fuel nozzles 114 (only one
illustrated) injects a steady stream of fuel into a combustor 132
that is made up of at least a combustor liner 134. The injected
fuel is ignited by a burner (not shown), which significantly
increases the energy of the high-pressure compressed air in the
combustor 132. This high-energy compressed air then flows first
into a high pressure turbine 115 and then a low pressure turbine
116, causing rotationally mounted turbine blades 118 on each
turbine 115, 116 to turn and generate energy.
[0015] The energy generated in the turbines 115, 116 is used to
power other portions of the engine 100, such as the fan module 110
and the compressor module 120. In particular, the turbines 115, 116
rotate a rotor 117 that extends through the engine 100, and the fan
module 110 and the compressor module 120 are mounted to the rotor
117. To contain the rotation of the rotor 117, one or more bearing
assemblies 119 are mounted around the rotor 117. The bearing
assembly 119 is attached to the remainder of the aircraft structure
via a bearing support housing 121.
[0016] The air exiting the combustor and turbine module 130 then
leaves the engine 100 via the exhaust module 140. The energy
remaining in the exhaust air aids the thrust generated by the air
flowing through the bypass 112.
[0017] After repeated use, one or more of the components of the
engine 100 may become damaged. In one example, the components may
be made of a sheet metal or alloy and may develop one or more
cracks due to either repeated thermal or mechanical stresses. These
components may include those that make up the combustor and turbine
module 130, such as the combustor liner 134, which may be
susceptible to cracking from exposure to excessive heat. In another
example, one or more components may be made of one or more alloys
having different properties, such as different thermal expansion
coefficients and mechanical properties. These components may also
have complex geometry shapes. Examples include components that act
as structural support for other engine components, such as bearing
support housings 121, which may have a complex shape and may be
made of different kinds of alloys. Turning to FIG. 2, a cross
section view of a component 200 including a damaged section 202 is
depicted, according to an embodiment. The damaged section 202 may
include a relatively small crack 204 that may measure between about
1 and 3 cm in length and between about 0.1 and 0.2 cm in depth.
[0018] In any case, these components may be difficult to repair
using conventional welding methods. In this regard, a method 300
depicted in a flow diagram in FIG. 3 may be used for repairing
these types of components. The method 300 includes removing surface
contaminants from an area around the crack 204 in a damaged section
202 of the component 200, step 302. Next, a braze paste is applied
to the damaged section 202 of the component 200, the braze paste
including a braze material, step 304. The damaged section 202 of
the component 200 is subjected to a first temperature that is below
a brazing temperature of the braze material to thereby decompose
the organic binder, step 306. The braze material is then heated to
a second temperature using laser energy, where the second
temperature is substantially equal to or above a brazing
temperature of the braze material to form the brazed joint, step
308. Each of these steps will now be discussed in more detail
below.
[0019] As mentioned above, contaminants may be removed from an area
around the crack 204, step 302. For example, contaminants, such as
oxides, may be chemically or mechanically removed from the
component. In an embodiment, the damaged section 202 and/or the
crack 204 may be subjected to vapor phase fluoride ion cleaning. In
such case, the component may be disposed in a container in which
fluoride ions, such as those in the form of hydrogen fluoride
vapor, are flowed over the component to remove oxides. The
component may then be subjected to a vacuum cleaning which may
substantially remove any remaining chemicals thereon. In another
embodiment, SiC carbide stones or metal cutters may be used to
physically remove contaminants from the damaged section 202 and/or
crack 204. Subsequently, the damaged section 202 and/or crack 204
may be cleaned with an acetone rinse.
[0020] Next, a braze paste may be applied to the damaged section
202, step 304. In an embodiment, the braze paste may be disposed at
least in the crack 204. In another embodiment, the braze paste
additionally may be applied to the area around the crack 204. The
braze paste may be made of a braze material and an organic binder.
It will be appreciated that the formulation of the braze material
may be tailored to a particular composition of the component. For
example, if the component is formed from a nickel-based superalloy,
the braze material may have a chemical composition that is
substantially similar to the nickel-based superalloy of the
component. In another embodiment, the component may be formed from
a cobalt-based superalloy; thus, cobalt-based braze materials
should be used to repair the defects. In any case, the braze
material may have a lower melting point than the component
alloys.
[0021] In an embodiment, the braze material includes a braze alloy
powder. The braze alloy powder may be any one of numerous metal or
alloy powders suitable for use in forming a brazed joint. For
example, in an embodiment, the braze alloy powder may be a powder
mixture that includes a high-melting temperature alloy powder and a
low-melting temperature alloy powder. A "high-melting temperature
alloy powder" may be defined as an alloy powder having a melting
temperature of between about 1260.degree. C. and about 1370.degree.
C. (e.g., about 2300.degree. F. and about 2500.degree. F.). A
"low-melting temperature alloy powder" may be defined as an alloy
powder having a melting temperature below 1150.degree. C. (e.g.,
about 2100.degree. F.), in an embodiment, or between 1065.degree.
C. and 1150.degree. C. (e.g., 1950.degree. F. and about
2100.degree. F.) in another embodiment, or as low as about
980.degree. C. (e.g., about 1800.degree. F.), in still another
embodiment.
[0022] Broadly, in an embodiment, the high-melting temperature
alloy powder may refer to an alloy powder that has a similar
composition to that of a nickel- or cobalt-based superalloy of a
gas turbine engine component to be repaired. In an embodiment of a
nickel-based high-melting temperature alloy powder, the powder may
include, by weight, about 60% Ni, about 10% W, about 10% Co, about
8.3% Cr, about 5.5% Al, about 1% Ti, about 3% Ta, about 0.1% Zr,
about 0.7% Mo, about 0.15% C, about 0.01% B and about 1.5% Hf. In
an embodiment of a cobalt-based high-melting temperature alloy
powder, the powder may include, by weight, about 10% Ni, about 7%
W, about 55% Co, about 23.5% Cr, about 0.2% Ti, about 3.5% Ta,
about 0.5% Zr, and about 0.6% C.
[0023] The low-melting temperature alloy powder may refer to an
alloy powder that includes a melting point depressant, such as
boron and/or silicon. In general, low-melting temperature alloy
powder has lower melting temperature than material from which the
component is made. The low-melting temperature alloy powder may
include, by weight, about 68.0% Ni, about 4.0% Al, about 3.5% Ta,
about 4% W, about 10.0% Co, about 9.0.0% Cr, 1.5% Hf and about 2.5%
B. In still another embodiment, the low-melting temperature alloy
powder may include, by weight, about 10% Ni, about 7% W, about
51.5% Co, about 23.4% Cr, about 0.2% Ti, about 0.5% Zr, about 0.6%
C, about 2.7% B, and about 0.4% Si.
[0024] The high-melting and low-melting temperature alloy powders
may be combined in a predetermined ratio to form the braze alloy
powder. The predetermined ratio may depend on the particular
material of the component to be repaired, the application for which
the component to be repaired will be used, the thermal environment
to which the component will be exposed, and other similar factors.
For example, the braze alloy powder may include a greater
percentage by weight of the high-melting temperature alloy powder
(e.g., greater than about 60%) if the component is to be subjected
to dimension and contour restoration in a later step. In an
embodiment, the braze alloy powder includes between about 40-70% of
the high-melting temperature alloy powder and between about 30-60%
of the low-melting temperature alloy powder. As alluded to above,
the braze alloy powder may be mixed with an organic binder. To form
the paste, the braze alloy powder and organic binder may be mixed
with a ratio of about 88 to about 12, by weight percentage.
[0025] The braze paste may be applied to the crack using any
suitable technique. In an embodiment, a paintbrush may be used to
paint the braze paste onto the component. In another embodiment,
the braze paste may have a relatively thin consistency and may be
poured onto the component. In still another embodiment, the braze
paste may be applied onto component using syringe
[0026] After the braze paste is applied at least to the crack 204,
the damaged section 202 is then subjected to a temperature suitable
to substantially decompose and evaporate the organic binder in the
paste, step 306. The phrase "substantially decompose" may be
defined as altering a microstructure of the organic binder such
that substantially all of the organic binders burn off. In an
embodiment, the component is heat-treated using a predetermined
temperature for a predetermined duration of time. For example, the
component may be disposed in a conventional vacuum furnace and
subjected to the predetermined temperature for the predetermined
duration of time. In another embodiment, the heat treatment may be
localized to the damaged section 202. For instance, a heating
apparatus, such as a laser system or hand-held laser, may be used
to heat the damaged section 202 of the component. The predetermined
temperature may be a temperature that is below the brazing
temperature (e.g. more than 600 degrees C. below) and at or above a
temperature at which the organic binder in the braze paste will
decompose or burn off. In particular, the predetermined temperature
is below a temperature at which the microstructure of the component
could not be altered. In an embodiment, the predetermined
temperature may be less than half the brazing temperature. In one
example, the braze material may have a brazing temperature of
1200.degree. C. and the predetermined temperature may be between
about 500.degree. C. and 550.degree. C., and preferably about
538.degree. C. The predetermined duration of time material may be a
duration that allows the organic binder to decompose and evaporate.
In an embodiment, the predetermined duration of time may be about 1
hour. It will be appreciated that the lower the temperature, the
more time may be employed, and vice versa.
[0027] The braze material is then heated using laser energy to a
second temperature that is substantially equal to or above the
brazing temperature to form the brazed joint on the component, step
308. The braze material may be directly heated or indirectly heated
with laser energy. The laser energy may be provided by a hand-held
laser. In an embodiment, the damaged section 202 and braze material
are subjected to a laser-welding process in which the laser energy
heats the braze material to a temperature substantially equal to or
above that of the high-melting temperature powder alloy therein. In
another embodiment, the damaged section 202 and braze material may
be subjected to a laser-brazing process. In laser-brazing, the
damaged section 202 is heated to the brazing temperature with a
laser, and the heat is conducted through the component and to the
braze material. Thus, the braze material melts without being
directly heated by the laser. To prevent contaminants from being
included in the resulting brazed joint, this step may occur in a
protective atmosphere. For example, the protective atmosphere may
be provided in a purge box that includes an inert gas, such as
argon, disposed therein.
[0028] In an embodiment, one or more post-brazing steps may be
performed, step 310. For example, the component may be machined to
an original shape and/or original dimensions. In another example,
at least one inspection process can be performed to determine
whether any surface defects, such as cracks or other openings,
exist. The inspection process can be conducted using any well-known
non-destructive inspection techniques including, but not limited
to, a fluorescent penetration inspection, and a radiographic
inspection. If an inspection process indicates that a component is
suitably in-situ braze-repaired, and then the repaired component is
ready for use.
[0029] The following example is presented in order to provide a
more complete understanding of the repair method 300. The specific
techniques, conditions, materials and reported data set forth as
illustrations, are exemplary, and should not be construed as
limiting the scope of the inventive subject matter.
[0030] In an example, a base metal substrate including a crack
thereon and made of Stellite.RTM. 31 superalloy supplied by
Stellite Coatings of Goshen, Ind. was subjected to a pre-braze
cleaning process. The cleaning processes included both hydrogen
fluoride ion cleaning and mechanical removal of oxides. A braze
paste was applied to the crack of the cleaned substrate. The braze
paste was made up of a mixture of a braze alloy powder and an
organic binder. The braze alloy powder included 50% by weight of a
high-melting temperature alloy powder, and 50% by weight of a
low-melting temperature alloy powder. The high-melting temperature
alloy powder included, by weight, 54.5% Co, 10.0% Ni, 23.5% Cr,
7.0% W, 3.5% Ta, 0.2% Ti, 0.60% C, and 0.5% Zr. The low-melting
temperature alloy powder included, by weight, 54.5% Co, 10.0% Ni,
23.5% Cr, 7.0% W, 3.5% Ta, 0.2% Ti, 0.60% C, 2.7%, 0.5% Zr, and
2.7% B. The organic binder was included in the braze alloy powder
at 12%, by weight.
[0031] The cleaned substrate was then exposed to a pre-braze heat
treatment at 538.degree. C. for an hour. The heat treatment was
used to substantially decompose the organic binder. Next, the braze
paste was laser-brazed with a hand-held laser set at 750 Watts
having a defocused laser beam of about 0.635 cm for between about
4-7 minutes to form a laser-brazed joint. The laser-brazed joint
dimension was about 2.54 cm in length, about 0.230 cm in width, and
about 0.152 cm in thickness. Optical photos showed the laser-brazed
joint to be metallurgically sound. Microhardness measurements were
taken of the laser-brazed joint and base alloy that fell between
HV300 and HV350, indicating that both the substrate and the brazed
joint had substantially similar microhardness properties. SEM
microphotographs indicated that elements making up the braze paste
and the substrate (except boron, which was not detected due to
equipment limitations) were uniformly distributed in both the
brazed joint and the substrate. Thus, by decomposing the organic
binder without altering the microstructure of the component and
before the step of brazing, a solid braze joint was formed.
[0032] Hence, an improved method for repairing cracks has been
provided. The method may repair the component without the formation
of additional cracks therein. In addition, the method may be
relatively simple and inexpensive to implement, as compared to
conventional repair methods. Moreover, by decomposing the organic
binder before the step of brazing, an improved brazed joint is
formed.
[0033] While at least one exemplary embodiment has been presented
in the foregoing detailed description of the inventive subject
matter, it should be appreciated that a vast number of variations
exist. It should also be appreciated that the exemplary embodiment
or exemplary embodiments are only examples, and are not intended to
limit the scope, applicability, or configuration of the inventive
subject matter in any way. Rather, the foregoing detailed
description will provide those skilled in the art with a convenient
road map for implementing an exemplary embodiment of the inventive
subject matter. It being understood that various changes may be
made in the function and arrangement of elements described in an
exemplary embodiment without departing from the scope of the
inventive subject matter as set forth in the appended claims.
* * * * *